(c)l999 American Institute of Aeronautics & Astronautics

AIAA-99-0037

NAVIER-STOKES SIMULATIONS OF THE NREL COMBINED EXPERIMENT PHASE II ROTOR

Earl P.N. Duque# ArmyLVASA Rotorcra$ Division

US Army Aeroflightrfynamics Directorate, Aviation Research, Development and Engineering Center

Aviation Missile Command NASA Ames Research Center

C. P. van Dam* Shannon C. Hughes** University of Cal$ornia, Davis

Abstract A Computational Fluid Dynamic method that solves the unstea& compressible Thin-layer Navier-Stokes equations was used to simulate the unste&y ji’owfield of the Combined &ueriment Phase II rotor Horizontal Axis Wind Turbine. 2%~ methodology makQs extensive use of overset grills to moaTeel the con&uration. Both isolated rotor and a complete conJguration that includes the rotor, hub, nacelle and tower were modeled The isolated rotor computations show good correlation with the available qetimental data. However, the low speeder had an adverse effect on solution convergence. In addition, the solutions demonstrate the importance of boundary layer transition The solutions for rotor-tower-nacelle configuration were also compared to available$eld data. 2%~ predicted unsteady normal force coeflcients show a tower wake interaction r:‘.^. that has a smaller duration and amplitude in comparison to the experiment. Although these first calculations point to some regions of the flowfield where the grid needs to be improved, they demonstrate a new capability that can simulate unsteae Wind Turbine aerodynamics.

Introduction The accurate prediction of wind turbine performance is critical during a machine’s development process. Performance predictions that are lower than the actual performance can lead to premature failure of the machine, increased maintenance costs and reduced revenue due to down time. In contra& greater than actual load predictions may reduce the failure rate and maintenance cost but result in a

# Research Scientist, ALW Member * Professor, Dept. of Mech. and Aero.

Engineering, AX4 Member * * Undergraduate Assistant,

Currently Lockheed-Martin - Sunnyvale.

This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States

heavier turbine and higher initial investment. This design is then less likely to produce electrical power at a competitive price.

Existing rotor models based on blade- element/momentum (BEM) theory can adequately predict rotor performance and loads. These models usually have restictions such as steady and uniform wind conditions. In addition, these methods make extensive use of two- dimensional airfoil data along with empirical models to determine stall effects, tower interference and three-dimensional effects such as tip relief.

These constraints do not accurately represent typical conditions encountered by either isolated turbines or turbines installed in wiud farms. For example, many horizontal-axis wind-turbines (HAWT) are designed to limit peak power output by stalling the blades. To maximize revenues, wind turbines tend to operate at or near maximum power output. Hence, these types of turbines frequently operate in stalled or near-

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stalled flow conditions. Stall is inherently unsteady and three-dimensional. These characteristics make it difficult to predict rotor performance.

Therefore, available models based on BEM theory continue to require improvements to be able to model rotor performance and airloads. These improvements can be derived from both experimental and computational methods. Through detailed experiments and computational methods, one can develop fundamental theories and methods to improve BEM simulations.

One such detailed experiment is the National Renewable Energy Lab (NREL) Combined Experiment shown in Figure 1. (Ref. 1. & 2) This turbine is a highly instrumented downwind, three-blade4 HAWT with rectangular blade plauform. Over the years, NREL has taken extensive surface pressure, flow visualization, meteorological and loads data from this field installed turbine. Furthermore, there are plans in place to test this turbine in NASA Ames Research Center’s National Full Scale Aerodynamic Center (NFAC) wind tunnel. Ibis test will allow for steady wind experiments and even more detailed aerodynamic measurements.

Figure 1 NRJXL Combined Experiment

NREL’s Combined Experiment Phase II HAWT is an instrumented three-bladed downwind ma&he that rotates at a constant 72 rpm and rated at 20 kW of electricrd power. This fixed- pitch rotor has a diameter of 33.25 ft., untwisted blades, and constant chord of 1.5 ft. The rotor uses the NREL S809 airfoil through out the span with some modifications towards the root to blend with the hub spar. The rotor hub is Exed at a pre-code of 3” and a fixed pitch of 12’. The rotor hub is mounted onto a nacelle that contains

the transmission and electrical power generation equipment. This assembly is then mounted 55ft from the ground on a cylindrical tower of 1.33 ft in diameter.

One of the three blades has flush mounted pressure taps and total pressure probes at several radial locations. In addition to the pressure data, NREL recorded the blade torque and bending loads as well as angle of attack at several radial stations. The inflow conditions for the turbine were measured by a vertical plane array of sensors positioned nominally upstream of the rotor. The details of NREL’s Combined Experiment are covered in Refs. 1 and 2..

This paper presents the results of a Computational Fluid Dynamic (CFD) study designed to support and complement the NFAC test. These computations solve the unsteady compressible Reynolds Average Navier-Stokes equations to obtain an accurate representation of the wind-turbine flowfield. This method has the ability to capture key unsteady flow features such as blade tip vortices, blade stall, the blade- tower wake interactions and sheared inflows. Both isolated rotor and rotor-tower interaction computations will be presented and demonstrate a capability to predict these difficult flotields. In addition, sensitivity to boundary layer transition to turbulence will be demonstrated and point to the need to either specify or measure transition in experiments.

Methodology The computer code used in this study is the OVERFLOW code version 1.6ap-rotorcraft by Buning (Ref. 3). I-his code solves the compressible Reynolds-Averaged Navier-Stokes (RaNS) equations using overset grids for the aerodynamic flowfield. Meakin (Ref. 4) tirst demonstrated the use of overset grids for rotor- body applications when he computed the flow-field of the V22 Osprey in forward flight. Ahmad & Duque (Ref. 5) extended Meakin’s method and demonstrated the aerodynamic capability of the method by computing the flowfield of an isolated two-bladed helicopter rotor in forward fight. The method successfully captured the unsteady wake development, downwash onto the rotor and included the specified longitudinal and lateral flap and cyclic blade motion. Recent developments by Duque et al. (Ref. 6) have extended the method to compute the farfield acoustic signature of arbitrary

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Figure 2: Blade Overset Surface Grids

helicopter rotors in forward flight and hover. In nacelle. All three blades and hub shanks were add&, Ahmad and Bauchau (Ref. 7) have modeled in both grid systems. In addition, the coupled a finite-element method to the grid system was designed to capture all the aerodynamics method and have shown a important geometric details such as the blade tip, favorable rotor dynamics prediction capability. the blade roof shank, and an idealized hub.

The OVERFLOW code has a number of solver options. In tbis study, the LU/SGS method by Jameson and Yoon (Ref. 8) was used to time advance the solution. This method has been shown to be unconditionally stable and when used with Newton-subiterations it can compute time-accurate solutions at time steps that correspond to % degree azimuth. In addition, 3ti order Roe’s scheme was employed for the RHS spatial terms.

The grids were generated such that the rotor spin axis lies along the Z coordinate axis, and the reference ground plane lies along a YZ plane that intersects some location in the negative X axis. The rotor rotates counter clockwise in the XY plane at Z=O with the wind direction along the negative Z axis.

Although a number of turbulence models are available in the OVERFLOW code including single equation and two-equation models, the Baldwin-Lomax (Ref. 9) model was chosen for its simplicity. Computations that assumed either fully turbulent or fkd transition were performed. The fully turbulent case assumed that the boundary layer is turbulent from the leading edge. For tied transition, the turbulence was enforced beginning at the % chord on both the upper and lower surfaces of the blade from root to the tip. The flow was assumed fully turbulent on the nacelle and on the tower. Although the iixed transition point is arbitrary, it does successfully demonskate the need to tix or document the transition in an experiment.

As shown in Figure 2, the grid system requires 5 overset grids per blade to properly defke the blade surface. These grids consist of a main blade grid, a tip cap grid, a grid to describe the blended surface at the hub shaft and then two more grids to define the juncture between the main rotor and the shaft. The hub and shafts require a total of three grids.

Structured grids were generated using the hyperbolic grid generator by Chan ( Ref. 10). These grids were formed such that their outer boundaries are no more than approximately half a chord away from the blade surface. Figure 3 illustrates typical body fitted volume grids for the hub-shank, blade and tip.

Two overset grids systems were generated for this study: isolated rotor and a rotor-tower-

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Figure 3 - Volume Grids

To complete the overset grid system, two stretched Cartesian grids surround the blade and hub volume grids. The first grid has uniform grid spacing in all three directions of approximately 0.05 of the chord. This first intermediate Cartesian grid stretches beyond the tip, above and below the minimum and maximum coordinates of the body fitted volume grids by a factor of 1.2 and a distance of 0.5 chord.

A second outer Cartesian grid covers the entire computational region. It initially has uniform spacing and surrounds the intermediate Cartesian grid. The grid space in this uniform region is approximately 0.1 chord. This grid then stretches by a factor of 1.2 and to a distance of 5 times the rotor radius to the outer boundary.

The rotor-tower-nacelle grid takes the key components of the isolated rotor grid system and

Figure 4: Nacelle and Tower Surface Grids

oversets them onto a tower-nacelle grid system. These grids include the tower, the nacelle and the rotor as described above. These grids were designed to capture the aerodynamic interference between the tower and rotor.

The surface and volume grids of the NREL combined experiment tower and nacelle was based upon drawings provided by NRJZL and from reference II! . The rotor and tower grid system consists of 5 additional surface grids. One grid covers the main tower while another covers the bulk of the nacelle body. Additional cap and collar grids were required to fully cover the ends of the nacelle and to join the nacelle and tower, respectively. Figure 4 highlights the current nacelle and tower surface grids. As in the isolated rotor case, the completed grid system is completed through the use of stretched Cartesian

Figure 5 - Cut Through all Grids and Tower Rotor Intersection

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grids.

The rotor and tower overset grid system makes use of the isolated rotor, hub and intermediate grid systems. The tower, nacelle and their related intermediate grids are simply overset onto the isolated rotor grid system A new outer Cartesian grid is then generated to model the ground plane and carry the solution to the far field. Figure 5 illustrates cutting plans that intersect through all the grids and highlights both the global Cartesian grids and also the interaction between the rotor and tower grid systems..

In total the overset grid system for the isolated rotor for the NFEL combine experiment consists of 2 1 individual overset grids. The grids vary in size from 100,000 for the smallest to 1.5 million for a grand total of 3.9 million grid points. The rotor-tower-nacelle grid system consists of 29 individual overset grids ranging from 100,OO for the smallest to 1.7 million points for the largest resulting in a grand total of 6.9 million points.

Results All the computations were performed on the

-12.00 r

-10.00 5 a)r/R=030

e270 degrees -T~UdtlOd -FIByTubtkd

Cray C-90 Systems at the NASA Ames Numerical Aerospace Simulation Facility. This system consists of a 16 processor, Cray C90 with a mtimum of 512 Megawords of main memory available to the user. In addition, the system has a large data storage subsystem that allows the user to rnaintanr and store the exorbitant amount of information generated.

All the solutions were ran for a total of 3-5 rotor revolutions. Each isolated rotor solution required approximately 16 single processor CPU-hours per revolution while the rotor-tower-nacelle computation required approximately 25 single processor CPU-hours per rotor revolution. Complete geomem and flow solutions were saved to permanent storage every 5 degrees of rotor azimuth totaling approximately 50 Gbytes per rotor revolution for the rotor-body-nacelle. The maximum m-core memory required is approtimately 150 Megawords.

All the solutions that follow correspond to flow conditions and experimental data provided by NREL. To facilitate and help verify the computations, flow conditions were chosen that

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0.40 0.60 0.80 1.00 1.20 0.00 0.20 0.40 0.60 0.80 1.00 120 x/Chord xlChord

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0.00 020 0.40 0.60 0.80 1.00 1.20 0.00 030 0.40 0.60 0.80 1.00 1.20 xlChord x/Chard

Figure 6: Pressure Coeffkient Comparison Between Experiment and Computation

U,,,=8.2m/s, 72RPM, RG, = l,OOO,OOO

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corresponds to “steady” wind conditions, 0’ nacelle yaw, and wind conditions that preclude blade stall. Table 1 lists the rotor flow conditions and the resulting flow-solver parameters.

Table 1 - Specified Experimental Flow Conditions and Corresponding Flow Sohrer

Parameters h 1

Experimental Conditions Rotation Rate 72 revolutions/minute Wind Velocity 8.22 meters/second Nacelle Yaw OD Reference Chord 0.4 meters Standard Altitude 5000 feet

Flow-Solver Parameters Tip Mach Number 0.11444 Tip Reynolds Number 1.02 Million Freestream Mach 0.0244 Blade Aspect Ratio 10.97 Time Sten 0.41825983 (l/4’?

Isolated Rotor The predicted surface pressure coefficients are compared to the NREL data at 4 spanwise locations- 30%, 47%, 63% and 80%. Figure 6 compares the experimental surface pressure coefficients to the tied-transition and fully turbulent case. These results provide some coniidence in the method’s ability to predict wind turbine rotor flowftelds. However, they also highlight the method’s inherent limitations.

The plots show the pressure distribution for four rotor azimuths at each radial location. An azimuth of 0”, shown by the diamond symbol, occurs when the rotor blade is oriented up and away from the support tower. The 90” position, shown by the square symbol, occurs before the blade encounters the tower. At 180” shown by the triangle, the blade points downward and lies directly behind the tower. Finally, the 270’ position, shown by the circle, occurs after the blade passes the rotor tower. The computed solutions were taken at the end of the last rotor revolution.

First, as expected, the experiment shows a strong dependence on rotor azimuth. The r/R+).33

section shows the greatest dependency on azimuth location. At O’, the blade exhibits a leading edge suction pressure coefficient of approximately -6.0. At 90°, the leading edge suction peak begins to reduce while at 180’ the upper surface pressure distribution flattens; indicating the presence of the support tower. After the blade passes the tower at 270’ the leading edge~suction peak begins to sharpen and increase in magnitude.

Since the computation is for an isolated rotor, we could assert that the solutions should correlate best with the 0” azimuth position. However, the r/R=O.33 surface pressures show a considerable discrepancy between the computations and the experiment. The fnst discrepancy is with the large overshoot with the leading edge stagnation pressure. Theoretically, the leading edge stagnation pressure should equal 1.0. However, the computed pressure overshoots to leading edge stagnation of approximately 2.

This error is consistent with known behavior for compressible formulations of the Navier-Stokes equations. It is well lmown that the compressible formulation of the Navier-Stokes will have convergence problems at freestream Mach numbers less than 0.2. The tip Mach number of this rotor is approximately 0.1144 while the freestream Mach number is approximate 0.0244. Therefore, a high Mach number of 0.1388 would be at the tip with the Mach number at inboard station of 0.33 equal to 0.037.

Although there is an error with the leading edge stagnation, the lower surface pressures past the leading edge are consistent with the experiment until the trailing edge. The trailing edge pressure tends to be under predicted indicating less trailing edge separation.

The calculations also show a strong dependency on turbulence transition as illustrated by the inboard stations of rEGO.33 and 0.47. The section at 0.33 shows the largest dependency. The leading edge suction peak reduces from a pressure coefficient of 11.0 for the fully turbulent case, down to levels comparable to the experiment for the ti transition case.

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x

4

Y

2

a) Perspective View

b) Side View C) Kear View I%cing upwind

Fire 7: Rotor-Tower Velocity Contours -Rotor at 0” Azimuth

Contour Levels: 0 to 0.05 iu 0.001 increments

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The outboard sections at r/R= 0.63 and r/R=080 show much more t&or-able comparisons and less dependency on transition. Both sections still show overshoots in the leading edge stagnation. However, the levels are lower than the inboard sections. The upper and lower surface pressures are consistent with the 0” azimuth experimental results. Furthermore, the transition cases tend to more closely follow the experiments than the fully turbulent results.

As with the inboard comparisons, the trailing edge pressures tend to be lower than the experiment. This inconsistency is both troublesome and revealing. The computations have some oscillations that may cause this error. In addition, the current turbulence model and ad hoc transition point can also lead to these errors.

However, we also suspect the inaccurate descriptions of the blade airfoil cross sections. For the computations, the airfoil cross sections are based upon published coordinates for the S809 airfoil. These coordinates are relatively coarse in comparison to the computational grid. Therefore, some interpolation was required to make a suitable surface de&&ion. In addition

and most importantly, the published coordinates describe a sharp-zero thiclmess trailing edge. This type of trailing edge is physically impossible to manufacture. Therefore, accurate measurements of the actual NREL Combined experiment rotor must be obtained and used in the computations.

Rotor-Tower Computations

Figure 7 illustrates a snapshot of the unsteady solution for the complete rotor-tower-nacelle configuration with the rotor at the 0” azimuth location after 3 rotor revolutions. It shows contours of velocity along cut planes of all the grids. The perspective view in Figure 7a highlights the overall features captured by the methodology. Two cut planes are shown: one near the ground plane and one that longitudinally bisects the machine along an XZ plane. The near ground plane highlights the presence of the tower wake. The tower wake appears rather uniform up until the vicinity of the main rotor.

The XZ plane in 7a) highlights the area where

-Experiment - Cak. Rotor-Tower - - -Cak. Isolated Rotor

Revolution Figure 8: Normal Force Coefficient Time EListory

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the rotor begins to have an effect on the tower flow field. Approximately halfway up from the ground plane there is a marked change where the contours turn inward toward the post and rotor plane. This change delineates the region where the rotor and tower begin to interact.

Figure 7b gives a side view of the XZ plane that bisects the machine. Again, we see a marked change in the contours as we follow them from the ground plane to about halfway up the tower. The contours also show the formation of the wake behind the rotor nacelle along with some contours that highlight the rotor wake. The side view brings out some of the grid density variations that can lead to loss of solution accuracy. In some of the areas, one can note multiple contour lines within grid overlap regions. The multiple lines result from disparate grid densities and aspect ratios. These differences can have an adverse effect on the convection of various flow features. For example, a separated flow region appears behind the nacelle and main rotor hub. About one nacelle length downstream, the resulting wake quickly dissipates indicating a change in grid density.

Figure 7c) illustrates the velocity flowfield along a YZ plane located one blade chord length behind the rotor hub spin axis. This cut plane further illustrates the formation of the main rotor wake interactions with the tower wake.

The unsteady normal and tangential forces and loads on the rotor blade are the most important quantities that the method should capture. Figure 8 compares the calculated normal force coefficient at the 80% radius location for both the isolated rotor and rotor-tower-nacelle cor@urations against the measured field loads. Both calculated loads assume fixed transition on the rotor. There is considerable agreement in the mean normal force coefficient between the solutions for the isolated and complete configurations.

The experimental data represents just 2 cycles of data. Both calculations overpredict the mean normal force. However, the calculations do show the effect of the tower wake at revolutions 1.5 and 2.5 ( 180” rotor azimuth ). In comparison to

the experiment, however, the computed wake interaction has neither the same duration nor amplitude as the experiment. In addition, the experiment exhibits a 2 per rev dominant frequency in normal force that needs further investigation.

Poor grid resolution and grid communication is the main cause of the disagreement in the tower wake and rotor interactions. Figure 9 illustrates the grids and contours of velocity on a YZ plane that corresponds to an 80% rotor radius when the rotor is directly behind the tower. The two sets were taken at a snapshot in time when the rotor blade was far from the tower and when the rotor was directly behind the tower (1 SO”).

We should first note tbat the wake behind the cylinder completely disappears at the grid boundary between the tower grid and the intermediate Cartesian mesh. At the overlap region, the solution has oscillations. Both of these clues indicate that we need better grid matching between the cylinder and Cartesian grid. In the second set of Figure 9, we can clearly see the airfoil interacting with the cylinder. However, upon closer inspection, the cylinder solution continues to have problems convecting its wake. Furthermore, the wake does not directly impinge upon the rotor blade.

Conclusions Tbis paper presents the results of a Navier- Stokes simulation of the NREL combined experiment Phase II rotor. The results give us insight into the rotor’s aerodynamic behavior. In particular, isolated rotor computations show a strong dependence on turbulence transition and on accurate trailing edge geometric definition. Complete rotor-tower-nacelle computations predict tower wake interactions. However, grid density mismatches had an adverse effect on the blade and wake interactions. Although some deficiencies exist, the results indicate that the method can predict the flowfield of an HAWT and demonstrate a significant increase in our ability to simulate Wind Turbine aerodynamics.

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2 --I

_ .-”

-..,,

Figure 9- Grid and Velocity Contours at r/R=O.SO

Contour Levels: 0 to 0.05 in 0.001 increments

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Acknowledgements This work was funded by the Department of Energy’s National Renewable Energy Lab through subcontract number DCX-7- 17211-o 1 entitiled “Army/NASA CFD Code Project”.

10. Chau W.M. & Meakin R.L., “Advances : Towards Automatic Surface Domain

Decomposition and Grid Generation for Overset Grids”, AIAA-97-1979,13” AIAA Computational Fluids Conference, Snowmass Village, Colorado, June 29- July

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